Aerodynamic static stability is defined as the tendency of the aircraft to return to an initial state or condition subsequent to or following a forced departure from such initial state or condition. Most commonly, aerodynamic static stability is operatively associated with the tendency of the aircraft to return, for example, to an initial angle of attack, angle of sideslip, or roll angle, subsequent to or following the occurrence of a particular disturbance.
The maneuvering capability of, for example, a fighter type aircraft is optimized if substantially neutral static stability is achieved and maintained. Excessive stability severely restricts the maneuvering capability of the aircraft, while excessive instability is highly undesirable in view of the fact that a highly-sensitive, fast-acting, and powerful automatic stability system must be provided for the aircraft and maintained functional at all times. Therefore, in order to cruise or maneuver most efficiently, an aircraft should be designed so that when trimmed, both the wings and the trimming surfaces are developing lift opposite in directional sense or orientation relative to the intended or impressed g loads such that the developed total lift, divided by the total drag, is a maximum. The wing-body or wing-fuselage assemblage, for example, should be designed to be substantially neutrally stable such that when experiencing both high and low load factors, the trimming aerodynamic surface is lightly loaded. In part, and in other words, there is reduced trim drag. This substantially neutral stability state can in fact be approximately achieved with current technology with respect to one particular flight condition and Mach number. It must also be appreciated at this juncture that the aircraft exhibits, or has operatively associated therewith, two different static stability components or characteristics. One of these static stability components or characteristics is lateraldirectional stability which operates about the roll and yaw axes, and the other static stability component or characteristic is longitudinal stability which operates about the pitch axis.
The problem now becomes quite apparent that with current technology, it is extremely difficult to design an aircraft that can in fact maneuver efficiently, that is, with a minimum amount of drag, throughout both the subsonic and high supersonic speed ranges, as well as within the transonic region. The reason for this is due to the drastically different pressure distributions or characteristics impressed upon the aircraft wings and control surfaces, fins, and stabilizers during the subsonic and supersonic speed regimes. Similar differences occur with respect to directional static stability as the aircraft experiences differing degrees of angle of attack.
A natural physical phenomenon of transonic aerodynamics, for example, is that longitudinal static stability increases with increasing Mach number, while directional static stability decreases with increasing supersonic Mach number. Similarly, directional static stability decreases with increasing angles of attack. The reasons for these phenomena are, for example, with respect to the change in the longitudinal stability of the aircraft, as the aircraft transcends or traverses the speed ranges from subsonic, through transonic, and into the supersonic ranges, the aerodynamic center shifts rearwardly from a position approximately 25% of the mean aerodynamic chord to one approximately 50% of the mean aerodynamic chord. This is due to the change in the pressure distribution over the aircraft airfoils from one having a substantially sharp pressure peak at the 25% mean aerodynamic chord position to a substantially uniform distribution over the airfoil.
With respect to the change in the directional static stability of the aircraft, the directional static stability of the aircraft is of course directly related to or dependent upon the lift curve slope of the vertical fin. It is well known that this slope changes with Mach number, and in fact peaks at Mach 1 and thereafter drastically drops off or sharply declines toward zero as the Mach number increases toward Mach 2 or beyond.
Similarly, with respect to the change in the directional static stability of the aircraft with increasing angle of attack, the vertical tail fin is normally disposed at a vertical location above the wing plane including the leading edge of the wing, and during normal, level cruise flight, this air region exhibits high energy air flow. Directional static stability is therefore high due to the interaction defined between the vertical fin/rudder and this high energy air flow. During high angles of attack, however, the vertical fin and rudder will be disposed below the normal wing plane as defined by the wing leading edge, and the energy of the air flow around the wing during such angles of attack has been substantially reduced due to the turbulent air flow conditions which prevail within the airfoil region when the airfoil is disposed at such angle of attack values. Consequently, the vertical fin and rudder cannot be as effective within such low energy air flow regions in order to impart to the aircraft the requisite directional static stability.
In view of the foregoing, it will therefore be readily appreciated that if an aircraft is designed to cruise and maneuver efficiently at subsonic speeds by being substantially neutrally longitudinally stable, then at supersonic speeds the aircraft will exhibit excessive longitudinal stability and will not be able to be maneuvered efficiently. Similarly, if the aircraft is designed to maneuver and cruise efficiently at supersonic speeds from a directional static stability point of view, then at subsonic speeds the aircraft will exhibit excessive directional static stability and will not be capable of efficient maneuverability.
In order to therefore practically accommodate the aforenoted incompatible flight characteristics of the particular aircraft which will transcend or traverse the subsonic, transonic, and supersonic speed ranges, several modes of design operations or practices have been currently instituted. For example, in order to accommodate the changes in longitudinal static stability of the aircraft with increasing Mach number, the aircraft has been specially designed with respect to the most important speed range for that aircraft and its designated missions, and the resulting penalties in performance and/or handling qualities or characteristics within the speed regions or regimes outside of the designed speed range must simply be tolerated. This means or mode of designing the aircraft and operating the same in light thereof is not at all ideal although such has been a conventionally acceptable solution to the stability and handling problems characteristic of supersonic aircraft. Another mode of operational design for supersonic aircraft in accommodating the changes in the longitudinal static stability of the aircraft with increasing Mach number is to design the particular aircraft such that the disposition or location of the center of gravity of the craft renders the aircraft slightly stable at supersonic speeds, the craft being capable of being efficiently trimmed and controlled. At subsonic speeds, the aircraft will be aerodynamically unstable. This condition or state is also not particularly desirable in view of the fact that an active control system is therefore required to impart the necessary stability to the aircraft. Such a system adds weight to the aircraft, and also adds a critical active element to the craft with the attendant risk of aircraft destruction should the control system experience failure.
A further current or conventional mode of operational design in conjunction with supersonic aircraft for accommodating the varying longitudinal static stability characteristics of the subsonic and supersonic speed regimes comprises the in-flight shifting of a substantially easily translatable massive entity within the aircraft, such as, for example, the fuel load or the like. The operational drawback or disadvantage with this type of stability design scheme resides in the fact that auxiliary apparatus must be provided for accomplishing the shifting of the massive entity, and therefore, the center of gravity of the aircraft. This apparatus adds considerable additional weight to the aircraft. In addition, such an operational scheme is not in fact desirable in view of the requirement that fighter aircraft frequently have to transcend or traverse the transonic speed range during the course of combat in an extremely rapid fashion. In a similar manner, in-flight variation of the wing geometry is also not particularly desirable in view of the requirement for massive actuators and increased structural weight attendant the movement of the wing structures. Fixed-geometry wing structures exhibiting highly swept, low aspect ratio characteristics and relatively small changes or variations in longitudinal stability at transonic speeds are likewise undesirable as such wings do not exhibit efficient stability characteristics at subsonic speeds.
A current or conventional mode of operational design for supersonic aircraft in order to accommodate the changes in the directional static stability of the aircraft as the craft transcends or traverses the subsonic, transonic, and supersonic speed ranges is to provide the aircraft with a vertical fin which has a geometrical size substantially greater than otherwise conventional subsonic aircraft vertical fins. In this manner, the directional static stability requirements for high supersonic speed flight are satisfied, however, the aircraft experiences excessive directional static stability during subsonic flight conditions. Alternatively, a smaller vertical fin can be employed for providing efficient directional static stability during subsonic flight conditions, however, an automatic stability system which controls the vertical tail fin or rudder as a function of sideslip is required.
Lastly, a current or conventional mode of operational design for supersonic aircraft in order to accommodate the large pitch down moment associated with vectored thrust in STOL operations is to provide the aircraft with an additional trim thrust device disposed forwardly of the center of gravity of the aircraft, and/or provide the aircraft with large and powerful aerodynamic control surfaces.
A need therefore exists for aircraft structure which can be incorporated into the structural systems or designs of supersonic aircraft so as to satisfy the various conflicting longitudinal and directional static stability requirements of the aircraft as the aircraft undergoes or experiences subsonic, transonic, and supersonic flight conditions, as well as varying degrees of angle of attack.
Accordingly, it is an object of the present invention to provide a new and improved aircraft structure which overcomes the aforenoted operational design disadvantages and drawbacks characteristic of conventional aircraft structure and structural systems.
Another object of the present invention is to provide a new and improved aircraft structure which is embodied within a new and improved empennage assembly.
Still another object of the present invention is to provide a new and improved empennage assembly for supersonic aircraft.
Yet another object of the present invention is to provide a new and improved empennage assembly for supersonic aircraft which will substantially improve the aerodynamic static stability of the aircraft throughout the subsonic, transonic, and high supersonic speed regimes and which can function as conventional pitch, yaw and roll controls.
Still yet another object of the present invention is to provide a new and improved empennage assembly for supersonic aircraft which will substantially improve the aerodynamic longitudinal and directional static stability components of the aircraft throughout the subsonic, transonic, and high supersonic speed regimes.
Yet still another object of the present invention is to provide a new and improved empennage assembly for supersonic aircraft which will substantially improve the aerodynamic static stability of the aircraft throughout the subsonic, transonic, and high supersonic speed regimes by rendering the aerodynamic longitudinal and directional static stability components of the aircraft substantially constant at an optimum low level of stability throughout the subsonic, transonic, and high supersonic speed ranges.
A further object of the present invention is to provide a new and improved empennage assembly for supersonic aircraft which substantially improves the maneuverability efficiency of the aircraft throughout the subsonic, transonic, and high supersonic speed ranges.
A still further object of the present invention is to provide a new and improved empennage assembly for supersonic aircraft which substantially improves the maneuverability capability and efficiency of the aircraft throughout the subsonic, transonic, and high supersonic speed regimes by rendering the aerodynamic static stability of the aircraft substantially constant independent of Mach number.
A yet further object of the present invention is to provide a new and improved empennage assembly for supersonic aircraft wherein the empennage assembly is of the caster type rotatable about a longitudinal axis disposed substantially parallel to the longitudinal axis of the aircraft.
A still yet further object of the present invention is to provide a new and improved empennage assembly for supersonic aircraft wherein the empennage assembly is of the caster type, comprising two tail surfaces of substantially different size, aspect ratio, sweep, and taper ratio, rotatable about a longitudinal axis disposed substantially parallel to the longitudinal axis of the aircraft so as to effectively interchange the angular orientation of the tail surfaces and thereby maximize the aerodynamic static stability characteristics of the aircraft throughout the subsonic, transonic, and high supersonic speed regimes.
A yet still further object of the present invention is to provide a new and improved empennage assembly for supersonic aircraft wherein the empennage assembly is of the rotatable caster type such that the angular orientation of the empennage is automatically achieved as a function of Mach number so as to maximize the aerodynamic static stability characteristics of the aircraft throughout the subsonic, transonic, and supersonic speed ranges.
An additional object of the present invention is to provide a new and improved empennage assembly for supersonic aircraft wherein the empennage assembly is of the rotatable caster type such that the angular orientation of the empennage can be varied in order to maximize the aerodynamic static stability efficiency characteristics of the aircraft depending upon, or as a function of, the varying degrees of angles of attack of the aircraft.
A still additional object of the present invention is to provide a new and improved empennage assembly for supersonic aircraft wherein the empennage assembly is of the rotatable caster type such that the angular orientation of the empennage can be varied in order to use wing downwash to help trim the pitching moments associated with vectored thrust, such as, for example, during STOL operations.